A long-recognized need in the turbine engine art has been to attain higher operating temperatures in order to achieve a greater thermodynamic efficiency and an increased power output per unit of engine weight. Ideally, a turbine engine should operate with stoichiometric combustion in order to extract the greatest possible energy value from the fuel consumed. However, the temperatures resulting from stoichiometric and even near-stoichiometric combustion are beyond the melting points of present-day metallic turbine engine components. Consequently, as the turbine engine art has progressed, an ever greater emphasis has been placed upon both cooling techniques and the development of temperature and oxidation resistant metals for use in components of the engine which are exposed to the highest temperatures. That is, cooling techniques and high temperature metals have been developed for each of combustion chambers, turbine stator nozzles, and turbine blades of turbine engines. This quest has led to the development of cooling schemes for all of these components as well as to classes of nickel-based "super alloy" metals which may be cast using directionally solidified or single crystal techniques. All in all, the quest for higher operating temperatures in a turbine engine fabricated of metallic components has led to increasing complexity and expense in the making of the engine, with ever-diminishing return for these efforts.
Unfortunately, as aircraft flight speeds have increased, the ability to use ambient air as a cooling mechanism has decreased. That is, the stagnation temperature of the ambient air goes up with increasing flight speed so that the liberation of unwanted heat to this air is difficult. Conventional engine cooling schemes, therefore, become less effective as flight speed increases. Also, compressor discharge air is used for turbine cooling. Higher performance engines have used higher pressure ratio compressors which in conjunction with high flight speeds causes this difficulty.
Particularly in the area of cooling the engine turbine stator nozzles and turbine blades, this increase of compressor discharge and ambient stagnation temperatures has presented difficulties. Because high-speed aircraft demand high engine performance, the cycle pressure ratio within the engine must be comparatively high. Thus, compressor discharge temperature of a high performance engine at high flight speed may be expected to reach about 1400.degree. F., or higher. Conventional wisdom, as evidenced by U.S. Pat. Nos. 3,528,250, issued 15 Sept. 1970, to D. Johnson: 4,254,618, issued 10 Mar. 1981, to Earnest Elovic: and 4,645,415, issued 24 Feb. 1987, to Edward J. Hovan, et al, teaches to reject heat from the pressurized compressor discharge air directly to ambient air, and then to use this somewhat cooled pressurized air for cooling within the engine. The limited effectiveness of this conventional cooling method for high flight speeds is recognized. For example, with compressor discharge temperature at 1400.degree. F. or higher, and an ambient-air heat exchanger effecting a temperature reduction of 100.degree. F. to 200.degree. F., the difficulty of cooling turbine components which are at 1600.degree. F. to 2000.degree. F. with a flow of air at 1200.degree. F. to 1300.degree. F. is self-evident. The volume of cooling air required becomes prohibitive.
A conventional alternative is presented by U.S. Pat. No. 3,083,532, issued 2 Apr. 1963, to H. Cook. This teaching adds the use of evaporative cooling with an expendable liquid, such as water. However, the disadvantage of having to carry a supply of water, or other liquid, and the consequences of depleting the liquid supply at a critical time mitigate against this proposed solution.
Yet another conventional teaching is provided by U.S. Pat. No. 3,314,649, issued 18 Apr. 1967, to J. R. Erwin, et al, wherein a peripherally outer portion of a turbine disk at the shank portion of the turbine blades defines a recycling or Terry-type turbine. Hot pressurized air obtained from ram effect or from a compressor is directed upon the recycling turbine and delivers work to the turbine disk. It is asserted that the gas-expansion cooling effect results in cooling of the turbine disk and turbine blades thereon. Unfortunately, the teaching of this patent also would add weight and structural complexity where added weight cannot be tolerated, and where all available structure must be devoted to turbine blade retention. This teaching has not been accepted in the field as a solution to turbine cooling.
An alternative approach to the attainment of higher operating temperatures in a turbine engine has been recognized. This approach involves the use of high-strength ceramic components in the engine. Ceramic components are generally better able than metals to withstand the high temperature oxidizing environment of a turbine engine. However, the term "high strength" in connection with ceramic structures must be viewed in context. While many ceramic materials exhibit superior high temperature strength and oxidation resistance, ceramics have historically been difficult to employ in turbine engines because of a comparatively low tensile fracture strength and a low defect tolerance. Consequently, ceramic structures have not widely replaced metallic components in the turbine engine field.